Note : Les descriptions sont présentées dans la langue officielle dans laquelle elles ont été soumises.
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GAS TURBINE ENGINE WITH AN OFFTAKE
FIELD OF THE INVENTION
[0001] The present invention relates to a gas turbine engine with an offtake.
BACKGROUND OF THE INVENTION
[0002] Turbine engines, and particularly gas or combustion turbine engines,
are rotary
engines that extract energy from a flow of combusted gases passing through the
engine
onto a multitude of rotating turbine blades. Typically a fan section is
utilized to take in
ambient air and direct it to different components of the engine for extracting
energy and
cooling purposes. Some of the fan air is initially directed into the
compressor stages, while
other portions of the fan air continue through outlet guide vanes and can
later be directed
into the engine components as needed.
[0003] Gas turbine engines include offtakes in areas of the engine where air
is extracted
from high-velocity, swirling channels to the internal air system for cooling,
sealing or heat
management purposes. When the angle of redirection is 90 or higher louvers or
other
aerodynamic shapes are required to turn the flow effectively. The louvers are
typically
cascades of equal length, shape and camber angle.
BRIEF DESCRIPTION OF THE INVENTION
[0004] In one aspect, embodiments of relate to gas turbine engine comprising
an annular
fan exhaust section, an engine core at least partially located within the fan
exhaust section,
a cooling air offtake located in the engine core and having an inlet, a louver
located at the
inlet and having at least two different size airfoils in spaced axial
arrangement.
[0005] In another aspect, embodiments relate to a louver assembly for an off
take of a
gas turbine engine comprising at least four airfoils in axial arrangement,
with none of the
airfoils are of the same size.
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BRIEF DESCRIPTION OF THE DRAWINGS
[0006] In the drawings:
[0007] Figure 1 is a schematic cross-sectional diagram of a gas turbine engine
for an
aircraft.
[0008] Figure 2 is an enlarged view of a fan exhaust section of the gas
turbine engine of
Figure 1.
[0009] Figure 3 is an enlarged view of an inlet with a louver having multiple
airfoils for
a cooling offtake duct for the gas turbine engine of Figure 1.
[0010] Figure 4A is a flow diagram of a conventional louver assembly.
[0011] Figure 4B is a flow diagram of an embodiment of the proposed louver
assembly.
DESCRIPTION OF EMBODIMENTS OF THE INVENTION
[0012] The described embodiments of the present invention are directed to a
gas turbine
engine have a louver to redirect fan air. For purposes of illustration, the
present invention
will be described with respect to the turbine for an aircraft gas turbine
engine. It will be
understood, however, that the invention is not so limited and may have general
applicability
within an engine, including compressors, as well as in non-aircraft
applications, such as
other mobile applications and non-mobile industrial, commercial, and
residential
applications.
[0013] As used herein, the term "forward" or "upstream" refers to moving in a
direction
toward the engine inlet, or a component being relatively closer to the engine
inlet as
compared to another component. The term "aft" or "downstream" used in
conjunction with
"forward" or "upstream" refers to a direction toward the rear or outlet of the
engine relative
to the engine centerline.
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[0014] Additionally, as used herein, the terms "radial" or "radially" refer to
a dimension
extending between a center longitudinal axis of the engine and an outer engine
circumference.
[0015] All directional references (e.g., radial, axial, proximal, distal,
upper, lower,
upward, downward, left, right, lateral, front, back, top, bottom, above,
below, vertical,
horizontal, clockwise, counterclockwise, upstream, downstream, aft, etc.) are
only used for
identification purposes to aid the reader's understanding of the present
invention, and do
not create limitations, particularly as to the position, orientation, or use
of the invention.
Connection references (e.g., attached, coupled, connected, and joined) are to
be construed
broadly and can include intermediate members between a collection of elements
and
relative movement between elements unless otherwise indicated. As such,
connection
references do not necessarily infer that two elements are directly connected
and in fixed
relation to one another. The exemplary drawings are for purposes of
illustration only and
the dimensions, positions, order and relative sizes reflected in the drawings
attached hereto
can vary.
[0016] Figure 1 is a schematic cross-sectional diagram of a gas turbine engine
10 for an
aircraft. The engine 10 has a generally longitudinally extending axis or
centerline 12
extending forward 14 to aft 16. The engine 10 includes, in downstream serial
flow
relationship, a fan section 18 including a fan 20, a compressor section 22
including a
booster or low pressure (LP) compressor 24 and a high pressure (HP) compressor
26, a
combustion section 28 including a combustor 30, a turbine section 32 including
a HP
turbine 34, and a LP turbine 36, and an exhaust section 38.
[0017] The fan section 18 includes a fan casing 40 surrounding the fan 20. The
fan 20
includes a plurality of fan blades 42 disposed radially about the centerline
12. The fan
casing 40 can also surround at least a portion of the fan exhaust section 41.
The HP
compressor 26, the combustor 30, and the HP turbine 34 form a core 44 of the
engine 10,
which generates combustion gases. The core 44 is surrounded by core casing 46,
which
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can be coupled with the fan casing 40, so that the core 44 is at least
partially located within
the fan exhaust section 41.
[0018] A HP shaft or spool 48 disposed coaxially about the centerline 12 of
the engine
drivingly connects the HP turbine 34 to the HP compressor 26. A LP shaft or
spool 50,
which is disposed coaxially about the centerline 12 of the engine 10 within
the larger
diameter annular HP spool 48, drivingly connects the LP turbine 36 to the LP
compressor
24 and fan 20.
[0019] The LP compressor 24 and the HP compressor 26 respectively include a
plurality
of compressor stages 52, 54, in which a set of compressor blades 56, 58 rotate
relative to a
corresponding set of static compressor vanes 60, 62 (also called a nozzle) to
compress or
pressurize the stream of fluid passing through the stage. In a single
compressor stage 52,
54, multiple compressor blades 56, 58 can be provided in a ring and can extend
radially
outwardly relative to the centerline 12, from a blade platform to a blade tip,
while the
corresponding static compressor vanes 60, 62 are positioned upstream of and
adjacent to
the rotating blades 56, 58. It is noted that the number of blades, vanes, and
compressor
stages shown in Figure 1 were selected for illustrative purposes only, and
that other
numbers are possible.
[0020] The blades 56, 58 for a stage of the compressor can be mounted to a
disk 59,
which is mounted to the corresponding one of the HP and LP spools 48,50, with
each stage
having its own disk 59, 61. The vanes 60, 62 for a stage of the compressor can
be mounted
to the core casing 46 in a circumferential arrangement.
[0021] The HP turbine 34 and the LP turbine 36 respectively include a
plurality of turbine
stages 64, 66, in which a set of turbine blades 68, 70 are rotated relative to
a corresponding
set of static turbine vanes 72, 74 (also called a nozzle) to extract energy
from the stream of
fluid passing through the stage. In a single turbine stage 64, 66, multiple
turbine vanes 72,
74 can be provided in a ring and can extend radially outwardly relative to the
centerline 12,
while the corresponding rotating blades 68, 70 are positioned downstream of
and adjacent
to the static turbine vanes 72, 74 and can also extend radially outwardly
relative to the
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centerline 12, from a blade platform to a blade tip. It is noted that the
number of blades,
vanes, and turbine stages shown in Figure 1 were selected for illustrative
purposes only,
and that other numbers are possible.
[0022] The blades 68, 70 for a stage of the turbine can be mounted to a disk
71, which is
mounted to the corresponding one of the HP and LP spools 48, 50, with each
stage having
its own disk 71, 73. The vanes 72, 74 for a stage of the compressor can be
mounted to the
core casing 46 in a circumferential arrangement.
[0023] The portions of the engine 10 mounted to and rotating with either or
both of the
spools 48, 50 are also referred to individually or collectively as a rotor 53.
The stationary
portions of the engine 10 including portions mounted to the core casing 46 are
also referred
to individually or collectively as a stator 63.
[0024] In operation, the airflow exiting the fan section 18 is split such that
a portion of
the airflow is channeled into the LP compressor 24, which then supplies
pressurized
ambient air 76 to the HP compressor 26, which further pressurizes the ambient
air. The
pressurized air 76 from the HP compressor 26 is mixed with fuel in the
combustor 30 and
ignited, thereby generating combustion gases. Some work is extracted from
these gases by
the HP turbine 34, which drives the HP compressor 26. The combustion gases are
discharged into the LP turbine 36, which extracts additional work to drive the
LP
compressor 24, and the exhaust gas is ultimately discharged from the engine 10
via the
exhaust section 38. The driving of the LP turbine 36 drives the LP spool 50 to
rotate the
fan 20 and the LP compressor 24.
[0025] A remaining portion of the airflow 78 bypasses the LP compressor 24
travelling
through the fan exhaust section 41 and exiting the engine assembly 10 through
a stationary
vane row, and more particularly an outlet guide vane assembly 80, comprising a
plurality
of airfoil guide vanes 82.. More specifically, a circumferential row of
radially extending
airfoil guide vanes 82 are utilized adjacent the fan section 18 to exert some
directional
control of the airflow 78. Upon exiting the fan exhaust section 41, the
airflow 78 can be
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redirected using a cooling air offtake 84 for additional cooling of the engine
core 44 and
turbine section 32.
[0026] Some of the ambient air supplied by the fan 20 can bypass the engine
core 44 and
be used for cooling of portions, especially hot portions, of the engine 10,
and/or used to
cool or power other aspects of the aircraft. In the context of a turbine
engine, the hot
portions of the engine are normally the combustor 30 and components downstream
of the
combustor 30, especially the turbine section 32, with the HP turbine 34 being
the hottest
portion as it is directly downstream of the combustion section 28. Other
sources of cooling
fluid can be, but is not limited to, fluid discharged from the LP compressor
24 or the HP
compressor 26. This fluid can be bleed air 77 which can include air drawn from
the LP or
HP compressors 24, 26 that bypasses the combustor 30 as cooling sources for
the turbine
section 32. This is a common engine configuration, not meant to be limiting.
[0027] Figure 2 is an enlarged view of the area near the fan exhaust section
41. The
cooling air offtake 84 comprises a duct 86 having walls 88 that turn through
nearly 900
from a primarily radial orientation to a primarily axial orientation. The
cooling air offtake
84 includes an inlet 90 located downstream of the outlet guide vane assembly
80. The inlet
90 includes a louver assembly 92 having a louver 93 comprising at least two
different size
airfoils 94, 96. While illustrated at a location downstream of the fan exhaust
section 41, the
offtake 84 can be located at any appropriate location throughout the engine.
[0028] In an exemplary embodiment illustrated in Figure 3 the louver assembly
92
includes four airfoils 94, 96, 98, 100 spaced in an axial arrangement. The
inlet 90 has a
leading edge 85 with a rounded lip and a trailing edge 87 having a chamfer
angle (3 of at
least 20 , but not to exceed 30 measuring from the duct wall 88 towards the
trailing edge
87 axially upstream. This feature will allow for higher pressure air bleed and
moving the
impingement point aft. An excessive angle will result in undesired pressure
losses.
[0029] The geometry of airfoil 96, which will be referred to as the primary
airfoil 96, is
outlined in Figure 4 described by a chord length C having a length defined as
a line from a
leading edge 108 to a trailing edge 110 and a height H having a length defined
as a line
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from a radial maximum 112 to a radial minimum 114 relative to the engine
centerline. Each
airfoil is also described by an angle of attack a measured from a local
relative wind
direction 116 to a continuous line along the chord length C. For illustrative
purposes the
dimensions for an initial, third, and fourth airfoil 94, 98, 100 will be
represented by
subscripts 1, 2, and 3 respectively.
[0030] The primary airfoil 96 is geometrically larger, both with respect to
the chord
length C and the height H, than the other three airfoils 94, 98, 100. The
maximum height
H of the primary airfoil 94 is at least 2 times larger than the maximum height
Hi of the
initial airfoil 94. The chord length C is at least 2.5 times larger than the
chord length Ci.
The axial arrangement of the airfoils comprises a geometry partially defined
by a chord
length relationship as follows:
C> C3 > > C4
[0031] The spaced axial arrangement includes the initial airfoil 94 nearest
the leading
edge 85 of the inlet 90, after which the primary airfoil 96 is located
downstream of the
initial airfoil 94, followed in the downstream direction by the third and
fourth airfoils 98,
100. The third and fourth airfoils 98, 100 are spaced equivalently so that the
distance
between the duct wall 88 and the fourth airfoil 100 is nearly the same as the
distance
between the third and fourth airfoils 98, 100. This spacing prevents flow
separation
between airfoils whilst keeping a Mach number high (See Figures 4A and 4B).
[0032] The angle of attack a for the third and fourth airfoils 98, 100 is
different than the
angle of attack a for the first and second airfoil 94, 96. In an exemplary
embodiment the
angle of attack a for the third and fourth airfoils 98, 100 is greater than
that of the first and
second airfoil 94, 96.
[0033] In an exemplary embodiment, the trailing edges 110 of the third and
fourth airfoils
98, 100 terminate in a line L connecting the trailing edge 110 of the primary
airfoil 96 to a
point 118 downstream of a trailing edge 87 of the inlet 90. This geometry
causes
corresponding chord lengths C3, C4 for the third and fourth airfoil 96, 98
become
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consecutively shorter. This relationship manages to turn effectively the flow
whilst
reducing any friction losses due to flow contact with the airfoil surface.
[0034] The overall benefit of the current embodiments is seen by the
comparison of Fig.
4A, showing a contemporary louver assembly with equal sized airfoils, with
substantially
the same angle of attack and equal spacing, as compared to the embodiment of
Figs. 2-3.
For the conventional louver assembly the flow direction is changed by guiding
the airflow
using a louver 122 having similar shaped airfoils 124 as depicted in Figure
4A. This design
can cause airflow separation 126 which is undesirable for effective airflow
movement.
Increasing the size of the primary airfoil 94 so that the louver assembly 92
comprises at
least two different size airfoils 94, 96 where the second 96 is geometrically
larger than the
first 94. This geometry differentiation causes an acceleration 128 of the flow
depicted in
Figure 4B allowing for a total engine pressure Pt increase.
[0035] Each of the airfoils 94, 96, 98, 100 in the louver assembly 92 is
designed with a
purpose, ensuring the effective use of the individual aerodynamic geometry.
The initial
airfoil 94 is configured to stabilize a boundary layer 130 and contain
recirculation 132 in
the duct 86. With a conventional louver assembly 120 the boundary layer 131 is
too thick
and will induce separation, wherein as seen in Figure 4B, the boundary layer
130 by both
the initial airfoil 94 and primary airfoil 96 is well defined. The primary
airfoil 96 is
configured to accelerate 128 the flow to maximum speed 134 without flow
separation. The
third and fourth airfoils 98, 100 are configured to guide the flow from
downstream of the
primary airfoil 96 in order to prevent separation.
[0036] Thorough CFD (Computational Fluid Dynamics) analyses has been conducted
and supports the benefit of the louver assembly 92 as compared to conventional
louver
assemblies 120. 2D optimization backed up with a 3D analysis has been carried
out with
tabulated results following. The pressure recovery is maximized whether
considering an
area from the fan exhaust section 41 to the HP turbine 34 or from the fan
exhaust section
41 to the LP turbine 36 both of which enable a reduction in bled flow. The
following table
compares a first engine recovery ratio to a second engine recovery ratio where
the second
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engine recovery ration includes the louver assembly 92 in place and the
pressure recovery
is at least 0.30. The goal is to maintain the highest total pressure (Pt) as
possible so as to
best move air through the duct to the turbine sections.
Total Pressure Static Pressure Total Pressure
Recovery = (Pt ¨ Ps13)/(P113¨ P913) Recovery
(Stage 13) Pt13 (Stage 13) Psi3
One Engine with conventional louver
0.131 8.07 6.82 6.99
design
Second Engine with proposed louver
0.362 8.369 7.097 7.557
assembly through the HP turbine
Second Engine with proposed louver
0.349 8.369 7.097 7.541
assembly through the LP turbine
[0037] Benefits to increasing the pressure recovery and reducing the mass flow
include
allowing for the duct flow to be reduced while maintain power. As the room for
designing
pipes is typically constrained, the introduction of this approach enables that
duct pipes to
be designed with more flexibility.
[0038] It should be appreciated that application of the disclosed design is
not limited to
turbine engines with fan and booster sections, but is applicable to turbojets
and turbo
engines as well.
[0039] While there have been described herein what are considered to be
preferred and
exemplary embodiments of the present invention, other modifications of these
embodiments falling within the scope of the invention described herein shall
be apparent
to those skilled in the art.
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